Cooling cascade test device for evaluating turbo-aerodynamic performance of super-high-altitude engine

Abstract

The application provides a cooling cascade test device for evaluating the turbine aerodynamic performance of an ultra-high-altitude engine, and relates to the field of turbine component design tests, and comprises: test blades, a plurality of test blades are installed side by side between two side cascade plates; a gas supply system is used for supplying cooling gas to the test blades; the gas supply system comprises a gas supply joint, a gas supply cavity and a pressure stabilizing cavity which are sequentially communicated; a plurality of gas supply joints are arranged on the top of the gas supply cavity, and a first flow regulating grid is arranged on the bottom of the gas supply cavity; the top of the pressure stabilizing cavity is connected with the bottom of the gas supply cavity, and a second flow regulating grid is arranged on the top of the pressure stabilizing cavity; a pressure measuring hole is arranged on the side wall of the pressure stabilizing cavity; the bottom of the pressure stabilizing cavity is connected with the cascade plate, so that the inner cavity of the pressure stabilizing cavity is communicated with the cooling channel inside the test blade; and a plurality of outflow grooves for cooling gas injection are arranged on the surface of the test blade along the span direction. The application can measure the inlet and outlet flow field parameters, the main flow of the wind tunnel and the cooling gas flow under a low Reynolds number environment, and can provide reliable data for the aerodynamic performance test of the planar cascade.

Description

Technical Field

[0001] This invention relates to the field of turbine component design testing and verification in the process of aero-engine design, and specifically to a cooling blade cascade test device for evaluating the aerodynamic performance of a high-altitude engine turbine. Background Technology

[0002] As one of the power units of an aero-engine, the turbine converts the thermal energy of the high-temperature combustion gas into the kinetic energy of the rotating blades through expansion and work, driving the upstream compressor and fan to rotate and maintain the engine's operating state. Based on the operating state of the turbine's basic components, aero-engine turbine components can be divided into high-pressure turbines and low-pressure turbines. Turbine blades, according to their operating temperature and environment, can be divided into two types: cooled blades requiring cooling air and uncooled blades. The operating temperature of the cooled blades is higher than the allowable temperature of their material; therefore, cooling channels need to be arranged inside the blades, and cooling is achieved through cooling air holes on the blade surface.

[0003] The aerodynamic performance of turbine blade cascades is one of the important factors affecting the overall performance of turbine components. However, due to the high cost and difficulty of testing turbine components, simplified tests are usually used to evaluate the performance indicators of turbine blades during the turbine blade design stage. Planar blade cascade testing is one of the simplified tests, used to obtain flow data of two-dimensional airfoils to guide blade design. Due to its advantages such as low test cost, easy testing, and fast result iteration, planar blade cascade testing is one of the most commonly used verification methods in the design process of aero-engine turbines.

[0004] However, obtaining accurate experimental data from planar blade cascade tests requires certain preconditions, namely, ensuring flow field homogeneity and flow duality. Flow duality is used to characterize the ability of flow to maintain two-dimensional flow on the radial cross-section. Accurate and reliable planar blade cascade tests should ensure that the measurement cross-section is not affected by the boundary layer of the side cascade plates or the flow in the corner region. For cooled blade cascades, the irregular mixing of cold air and mainstream flow caused by cold air injection significantly disrupts flow duality, reducing the reliability of planar blade cascade test results.

[0005] Under real-world operating conditions, when an aircraft flies at an altitude greater than 16 kilometers, the low Reynolds number flow caused by low density at high altitudes becomes significant, resulting in substantial flow shear losses and becoming a major factor restricting the high-performance aerodynamic design of turbine engines. Simultaneously, the Reynolds number is a crucial factor affecting flow field uniformity and flow duality, thus influencing the accuracy and reliability of planar blade cascade tests. However, under current research and testing conditions, constructing a high-altitude low Reynolds number flow environment is difficult, and testing equipment adapted to low Reynolds numbers is lacking. Regarding blade cascade tests, current planar blade cascade testing equipment is often designed for standard ground-based atmospheric environments, failing to consider the impact of large-scale separation and flow inhomogeneities caused by cooling gas mixing after the Reynolds number decreases. In particular, conventional planar blade cascade test specimens only test the aerodynamic performance of the blade profile, without considering the additional effects of air-cooled layouts, and are clearly unsuitable for testing air-cooled turbine blades under high-altitude conditions. Therefore, if a turbine planar blade cascade test piece that takes into account the influence of cold air and realizes secondary airflow supply can be designed, and at the same time maintains good flow duality index under low density conditions, it will be of great significance to accurately evaluate the effectiveness of turbine blade design and injection cooling structure design, especially turbine performance design under high-altitude extreme conditions. Summary of the Invention

[0006] In view of this, this application provides a cooling blade cascade test device for evaluating the aerodynamic performance of a high-altitude engine turbine. This device can be used in a low Reynolds number environment and takes into account the effects of low Reynolds number and the influence of the internal cold air passage and cold air injection on the flow field of the planar blade cascade. This solves the problems of uneven cooling air supply, poor flow duality, the inapplicability of high Reynolds number cooling planar blade cascades to low Reynolds number working environments, and the high cost, high control difficulty, and low accuracy of conventional flow duality control methods.

[0007] This application provides the following technical solution: a cooling blade cascade test device for evaluating the aerodynamic performance of a high-altitude engine turbine, comprising:

[0008] A test blade, wherein grid plates are provided on both sides of the test blade, and multiple test blades are fixedly installed side by side between the grid plates on both sides;

[0009] An air supply system is installed on the grid plates on both sides to provide cooling gas to the test blades. The air supply system includes an air supply connector, an air supply chamber, and a pressure stabilizing chamber connected in sequence. The top of the air supply chamber is provided with multiple air supply connectors, and the bottom is provided with a first rectifier grid. The top of the pressure stabilizing chamber is connected to the bottom of the air supply chamber, and the top of the pressure stabilizing chamber is provided with a second rectifier grid. The side wall of the pressure stabilizing chamber is provided with a pressure measuring hole.

[0010] The test blade is provided with a blade-shaped groove for positioning and mounting on the grid plate; the bottom of the pressure stabilizing cavity is connected to the grid plate, so that the inner cavity of the pressure stabilizing cavity is connected to the cooling channel inside the test blade; both ends of the test blade extend to the inner side of the grid plate and into the interior of the pressure stabilizing cavity; multiple outlet grooves for cold air injection are formed along the spanwise direction on the surface of the test blade, and the outlet grooves include at least one long groove located in the middle region of the blade and several short grooves located on both sides of the long groove.

[0011] According to one embodiment of this application, the number of gas supply connectors is four, and the four gas supply connectors are evenly arranged on the top of the gas supply chamber.

[0012] According to one embodiment of this application, the first rectifier grille and the second rectifier grille are honeycomb grilles or mesh grilles.

[0013] According to one embodiment of this application, the pressure measuring holes are symmetrically distributed on both sides of the pressure stabilizing cavity, and the total number is not less than 40.

[0014] According to one embodiment of this application, a blade fixing baffle is further included. The blade fixing baffle is installed on the outer side of the grid plates on both sides by fasteners for fixing and pressing the end of the test blade.

[0015] According to one embodiment of this application, the plurality of test blades include at least two measuring blades located in the middle and test liner blades located on both sides of the measuring blades; static pressure measuring holes are formed on the surface of the measuring blades.

[0016] According to one embodiment of this application, the static pressure measuring hole is L-shaped, formed by the intersection and connection of a channel perpendicular to the blade surface and a channel along the blade span, and is arranged alternately with the cooling channels inside the blade.

[0017] According to one embodiment of this application, the outflow groove specifically includes a long groove located in the middle region of the blade and six short grooves symmetrically distributed on both sides of the long groove.

[0018] According to one embodiment of this application, the air supply chamber and the pressure stabilizing chamber, as well as the pressure stabilizing chamber and the grid plate, are all connected by bolts, and a sealing gasket is provided between the connecting surfaces.

[0019] According to one embodiment of this application, the test device is installed on the angle-of-attack adjustment disk of the planar blade cascade wind tunnel inside the variable density test chamber.

[0020] Compared with traditional planar blade cascade test devices, the test device designed in this invention has the following main advantages:

[0021] 1) This invention realizes blade air supply and blade surface cold air injection on the basis of traditional planar blade test device, and has the advantages of uniform and adjustable air supply pressure of multiple rows of blades, which greatly reduces air supply airflow disturbance.

[0022] 2) This invention uses multiple air supply chambers combined with a multi-layer honeycomb structure to rectify the complex vortex structure in the air supply chambers, while obtaining a uniform pressure field in the air supply chambers, so that the pressure and velocity at the cold air inlet of all blades are consistent.

[0023] 3) This invention designs the distribution of cooling gas grooves on the blade surface and arranges pressure relief grooves to ensure that the test groove location can maintain a uniform two-dimensional outflow under extremely low Reynolds number conditions, reduce radial non-uniformity, and obtain accurate blade aerodynamic parameters.

[0024] 4) The experimental apparatus of the present invention can reduce the influence of the two side grids on the flow in the mid-section under low Reynolds number conditions by means of the jet of the vent hole without using a complex suction system, and play a role in controlling the duality of flow at low cost.

[0025] 5) This invention can measure inlet and outlet flow field parameters and wind tunnel mainstream and cold air flow in low Reynolds number environments, providing reliable data for planar blade aerodynamic performance testing. Attached Figure Description

[0026] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0027] Figure 1 This is a schematic diagram of the air-cooled planar blade cascade test device designed for this invention;

[0028] Figure 2a and Figure 2b These are schematic diagrams of the air supply chamber of the present invention; wherein Figure 2a This is a side view. Figure 2b It is a bottom view;

[0029] Figure 3a and Figure 3b These are schematic diagrams of the voltage stabilizing cavity of the present invention; wherein Figure 3a This is a side view. Figure 3b This is a top view;

[0030] Figure 4a and Figure 4b These are schematic diagrams illustrating the connection structure of the grid plate, blade fixing baffle, and test blade of the present invention; wherein... Figure 4a This is a side view. Figure 4b This is a top view;

[0031] Figure 5 This is a schematic diagram of the structure of the test blade of the present invention;

[0032] Figure 6 This is a schematic diagram of the arrangement of pressure measurement holes on the blade surface of the present invention;

[0033] Figure 7 This is a schematic diagram of the wind tunnel installation for the test device;

[0034] Figure 8 This is a schematic diagram of the gas supply status of the experimental device of the present invention;

[0035] Figure 1 In the middle, 1-air supply connector; 2-air supply chamber; 3-pressure stabilizing chamber; 4-grid plate; 5-test blade; 6-blade fixing baffle; 7-bolt and nut assembly; 8-first group of bolts; 9-second group of bolts; 10-pressure measuring hole;

[0036] Figure 7 In the middle, 7.1-wind tunnel duct; 7.2-turbulence intensity adjustment grid; 7.3-converging section; 7.4-angle of attack adjustment disk; 7.5-variable density test chamber; 7.6-ejector port; 7.7-planar blade cascade test device. Detailed Implementation

[0037] The embodiments of this application will now be described in detail with reference to the accompanying drawings.

[0038] The following specific examples illustrate the implementation of this application. Those skilled in the art can easily understand other advantages and effects of this application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of this application, and not all of them. This application can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of this application. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. Based on the embodiments in this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0039] This invention provides a cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine, comprising:

[0040] A test blade, wherein grid plates are provided on both sides of the test blade, and multiple test blades are fixedly installed side by side between the grid plates on both sides;

[0041] An air supply system is installed on the grid plates on both sides to provide cooling gas to the test blades. The air supply system includes an air supply connector, an air supply chamber, and a pressure stabilizing chamber connected in sequence. The top of the air supply chamber is provided with multiple air supply connectors, and the bottom is provided with a first rectifier grid. The top of the pressure stabilizing chamber is connected to the bottom of the air supply chamber, and the top of the pressure stabilizing chamber is provided with a second rectifier grid. The side wall of the pressure stabilizing chamber is provided with a pressure measuring hole.

[0042] The test blade is provided with a blade-shaped groove for positioning and mounting on the grid plate; the bottom of the pressure stabilizing cavity is connected to the grid plate, so that the inner cavity of the pressure stabilizing cavity is connected to the cooling channel inside the test blade; both ends of the test blade extend to the inner side of the grid plate and into the interior of the pressure stabilizing cavity; multiple outlet grooves for cold air injection are formed along the spanwise direction on the surface of the test blade, and the outlet grooves include at least one long groove located in the middle region of the blade and several short grooves located on both sides of the long groove.

[0043] The cooling blade test device of this invention can perform planar air-cooled blade performance testing in a low Reynolds number environment. Considering the influence of low Reynolds number effect and the effect of internal cold air passage and cold air injection on the flow field of the planar blade, the main problems solved by this invention are as follows: (1) The problem of difficulty in achieving uniform cooling air supply due to the limitations of planar blade structure and size and the flow differences of each row of blades; (2) The problem of poor flow duality caused by uneven cold air injection of air-cooled turbine and complex mixing of cold air with mainstream; (3) The problem that the molded blades of cooling planar blades under high Reynolds number cannot be applied to low Reynolds number working environment; (4) Conventional flow duality control methods are often accompanied by complex test piece design and suction module, which are costly, difficult to control and have low accuracy.

[0044] like Figure 1 The diagram shows the planar blade cascade test device designed in this embodiment of the invention. The device consists of an air supply connector 1, an air supply chamber 2, a pressure stabilizing chamber 3, a cascade plate 4, test blades 5, a blade fixing baffle 6, and bolt and nut installation accessories (bolt and nut assembly 7, first set of bolts 8, and second set of bolts 9). The pressure stabilizing chamber 3 has pressure measurement holes 10 on its surface for monitoring the cold air inlet pressure. The test device has a symmetrical layout and uses sealing gaskets to ensure flow sealing in each chamber. However, the wall surface may leak air under low Reynolds number conditions.

[0045] In specific implementation, such as Figure 2a and Figure 2bThe diagram shows the air supply chamber in the device of the present invention. The air supply chamber has four air supply ports at the top and a mesh grid at the bottom. The air supply chamber is connected to the pressure stabilizing chamber by bolts and nuts. Its function is to collect four streams of cold air, so that the uneven airflow is fully mixed inside the air supply chamber, and then flows out through the honeycomb grid. The honeycomb grid can be used to break up large reflux vortex structures and obtain a more uniform outlet airflow pressure.

[0046] like Figure 3a and Figure 3b The diagram shows the pressure stabilizing chamber in the device of this invention. The top of the pressure stabilizing chamber is connected to the air supply chamber, which is also equipped with a mesh grid, forming a double-layer honeycomb structure with the air supply chamber. This structure effectively breaks up large backflow vortices and stabilizes the airflow. The bottom is connected to the grid plate via nuts. The airflow within the pressure stabilizing chamber is further adjusted, resulting in a slower flow rate and more uniform pressure. Pressure measuring holes are arranged on the surface of the pressure stabilizing chamber to detect the internal pressure and to calculate the cold air inlet parameters. There are 20 pressure measuring holes on each side of the pressure stabilizing chamber, for a total of 40 points, which can be used to provide feedback on the pressure uniformity within the chamber.

[0047] like Figure 4a and Figure 4b The diagram shows the assembly of the grid plates, blade fixing baffles, and test blades in the device of this invention. The grid plates have airfoil grooves, and the test blades are mounted on the grid plates on both sides. The airfoil grooves on the grid plates fix the installation angle of the blades, serving a positioning function. Under low Reynolds number conditions, influenced by unsteady flow aerodynamic forces and the internal air supply of the blades, the test blades may exhibit cross-movement, causing a shift in the test section. Therefore, radial fixing of the blades is necessary, hence the blade fixing baffles. These baffles are bolted to the grid plates on both sides, pressing the blades together during the test to provide fixation while not obstructing the air inlet cross-sectional area of ​​the air-cooled blades. The blade fixing baffles effectively prevent blade cross-movement, ensuring the positional accuracy of the measurement section. The entire device structure is optimized for the strong viscosity and easy separation characteristics of low Reynolds number flows.

[0048] like Figure 5The image shows the test blades in the device of this invention. The test device contains a total of eight blades, of which blades 1-3 and 6-8 are test support blades used to construct a stable, uniform intermediate flow field with good periodicity, and blades 4 and 5 are test blades, corresponding to an intermediate flow field with good periodicity. During the experiment, static pressure measurement holes and outlet flow field probes were arranged at the exit points of blades 4 and 5. After periodic verification, the aerodynamic performance of the blade cascade was measured. Static pressure holes were also opened on the pressure surface of blade 4 and the suction surface of blade 5 to detect the pressure distribution on the blade surface. Considering the blade air supply under low Reynolds number conditions, the length of all test blades was longer than the distance between the two side cascade plates. The blade coolant inlet was extended deep into the pressure stabilizing chamber to reduce the influence of the cascade wall on the coolant inlet flow. To ensure uniformity of coolant outflow, unlike the conventional film cooling method for air-cooled blades, film cooling channels were used to improve the radial uniformity of the coolant after mixing with the mainstream. Furthermore, to reduce the radial velocity variation of the coolant outflow caused by pressure and velocity attenuation along the blade's surface, in addition to the long groove in the central region as the test area, six short grooves were symmetrically arranged on both sides of the blade surface as drainage areas. Profile pressure measurement holes were separately arranged on blades 4 and 5. Figure 6 The diagram shows the mid-section structure and the corresponding axial position of the blade. The hydrostatic measurement holes on the profile are staggered with the blade cooling channels, forming an overall L-shape. This is achieved by connecting two pressure-inducing holes that are perpendicular to the blade surface and along the blade span.

[0049] The experimental setup operates in a low Reynolds number flow environment, characterized by strong viscosity and large separation scale. This results in an increased influence of endwall viscosity on the mainstream during planar blade cascade experiments, making it difficult to adjust the flow duality. The experimental setup targets a cooled blade structure that requires cold gas injection from the blade surface to capture mixed aerodynamic properties, which presents challenges such as narrow internal air supply channels and uneven cold gas outflow.

[0050] From the perspective of the experimental setup's structure, apart from the air supply interface, air supply chamber, pressure stabilizing chamber, and blade fixing baffle, the blades and cascade plates are the core components of a conventional planar blade cascade experimental setup. Therefore, this experimental setup also incorporates the testing functions of a conventional planar blade cascade test. Furthermore, this experimental setup proposes targeted solutions to the two major problems mentioned above.

[0051] During use, the planar blade cascade test apparatus of the present invention is installed on a variable-density planar blade cascade test bench. Figure 7The installation diagram shows that the grating plate of the planar blade cascade test device 7.7 is connected to the angle-of-attack adjustment disk 7.4 of the planar blade cascade test stand via an assembly. During the test, the airflow angle of the incoming flow in the wind tunnel duct 7.1 is adjusted by rotating the disk to obtain the inlet angle of attack of the corresponding test blade cascade. The incoming flow from the wind tunnel passes through the turbulence adjustment grid 7.2 for inlet turbulence adjustment, and then enters the test device through the convergence section 7.3. The variable density test chamber 7.5 realizes the low Reynolds number blade cascade test through the ejector port 7.6.

[0052] like Figure 8 As shown, during the test, an external pipe connects to the air supply port on the air supply chamber. Cool air enters through the air supply port, passes through the air supply chamber (first-layer rectification), and the pressure stabilizing chamber (second-layer rectification and pressure stabilization), before entering the cooling channel inside the blade. Finally, it flows out from the blade surface and mixes with the mainstream flow in the wind tunnel before flowing downstream. A mass flow meter should be installed on the air supply pipe to detect the flow rate of the cool air entering the blade. During the test, the mainstream flow rate in the blade channel is obtained by measuring the inlet parameters of the test section, while the cool air flow rate is obtained by the mass flow meter on the air supply pipe. By measuring the wind tunnel inlet parameters, the pressure stabilizing chamber pressure, the cool air flow rate, and the pressure on the blade surface and at the outlet, the aerodynamic performance of the cooling blade under low Reynolds number conditions can be comprehensively evaluated.

[0053] After entering the air supply chamber, the cold air is in a turbulent and disordered state with significant pressure unevenness. Passing through two layers of honeycomb holes, the turbulent airflow is rectified and enters the pressure stabilizing chamber, achieving a more uniform pressure distribution within the chamber and ensuring good uniformity of the airflow entering each blade. Because the pressure within the pressure stabilizing chamber is uniform, the airflow parameters entering the blade can be obtained through pressure measurement and control within the pressure stabilizing chamber. Unlike traditional film cooling orifices, this invention features multiple outflow slots along the spanwise direction on the blade surface. Specifically, a long slot is created in the central region of the blade (corresponding to the region with optimal flow duality) as the main testing area; six short slots are symmetrically created on both sides of the long slot as drainage areas. This "six short and one long" slot layout facilitates obtaining a more stable and better dual-uniform airflow at the long slot in the middle of the blade. This is because the radial velocity and pulsating disturbances of the airflow are pre-released and homogenized by the upstream short slots, ensuring that the cold air flowing out from the central long slot has excellent radial uniformity and two-dimensionality.

[0054] The experimental apparatus of this invention, through the installation of an air supply system with a double-layered rectifier grid (air supply chamber and pressure stabilizing chamber), can fully mix, rectify, and stabilize the turbulent cold air input from multiple sources. This ensures highly uniform cold air pressure and flow rate entering the cooling channel of each test blade, reducing the disturbance of the mainstream flow field caused by uneven cold air supply at the source and achieving uniform cold air supply under low Reynolds number conditions. Furthermore, the test blades employ an outflow slot design with a long slot (test area) and short slots on both sides (drainage areas), significantly improving the flow duality under low Reynolds number conditions. This invention enables the measurement of inlet and outlet flow field parameters and the mainstream flow and cold air flow rate in wind tunnels under low Reynolds number environments, providing reliable data for the aerodynamic performance testing of planar blade cascades.

[0055] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine, characterized in that, include: A test blade, wherein grid plates are provided on both sides of the test blade, and multiple test blades are fixedly installed side by side between the grid plates on both sides; An air supply system is installed on the grid plates on both sides to provide cooling gas to the test blades. The air supply system includes an air supply connector, an air supply chamber, and a pressure stabilizing chamber connected in sequence. The top of the air supply chamber is provided with multiple air supply connectors, and the bottom is provided with a first rectifier grid. The top of the pressure stabilizing chamber is connected to the bottom of the air supply chamber, and the top of the pressure stabilizing chamber is provided with a second rectifier grid. The side wall of the pressure stabilizing chamber is provided with a pressure measuring hole. The test blade is provided with a blade-shaped groove for positioning and mounting on the grid plate; the bottom of the pressure stabilizing cavity is connected to the grid plate, so that the inner cavity of the pressure stabilizing cavity is connected to the cooling channel inside the test blade; both ends of the test blade extend to the inner side of the grid plate and into the interior of the pressure stabilizing cavity; multiple outlet grooves for cold air injection are formed along the spanwise direction on the surface of the test blade, the outlet grooves including at least one long groove located in the middle region of the blade and several short grooves located on both sides of the long groove; The plurality of test blades includes at least two measuring blades located in the middle and test liner blades located on both sides of the measuring blades; static pressure measuring holes are formed on the surface of the measuring blades; The static pressure measuring hole is L-shaped, formed by the intersection of channels perpendicular to the blade surface and channels along the blade span, and is arranged alternately with the cooling channels inside the blade.

2. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The number of gas supply connectors is four, and the four gas supply connectors are evenly arranged on the top of the gas supply chamber.

3. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The first and second rectifier grilles are honeycomb grilles or mesh grilles.

4. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The pressure measuring holes are symmetrically distributed on both sides of the pressure stabilizing chamber, and the total number is not less than 40.

5. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, It also includes blade fixing baffles, which are installed on the outer side of the grid plates on both sides by fasteners to fix and press the ends of the test blades.

6. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The outflow groove specifically includes a long groove located in the middle region of the blade and six short grooves symmetrically distributed on both sides of the long groove.

7. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The air supply chamber and the pressure stabilizing chamber, as well as the pressure stabilizing chamber and the grid plate, are all connected by bolts, and sealing gaskets are provided between the connection surfaces.

8. The cooling blade cascade test apparatus for evaluating the aerodynamic performance of a high-altitude engine turbine according to claim 1, characterized in that, The test apparatus is installed on the angle-of-attack adjustment disk of the planar blade cascade wind tunnel inside the variable density test chamber.

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